JP4778754B2 - Cooling system for trailing edge of turbine bucket airfoil - Google Patents

Cooling system for trailing edge of turbine bucket airfoil Download PDF

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JP4778754B2
JP4778754B2 JP2005260354A JP2005260354A JP4778754B2 JP 4778754 B2 JP4778754 B2 JP 4778754B2 JP 2005260354 A JP2005260354 A JP 2005260354A JP 2005260354 A JP2005260354 A JP 2005260354A JP 4778754 B2 JP4778754 B2 JP 4778754B2
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airfoil
cooling
trailing edge
along
tip
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JP2006083851A (en
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アリエル・シーザー・プレペーナ・ジャカラ
ゲーリー・エム・イツェル
ジョシュア・アール・コーナウ
アザダリ・エイ・ラドハニ
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General Electric Co
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General Electric Co
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    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
    • F01D5/12—Blades
    • F01D5/22—Blade-to-blade connections, e.g. for damping vibrations
    • F01D5/225—Blade-to-blade connections, e.g. for damping vibrations by shrouding
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
    • F01D5/12—Blades
    • F01D5/14—Form or construction
    • F01D5/18—Hollow blades, i.e. blades with cooling or heating channels or cavities; Heating, heat-insulating or cooling means on blades
    • F01D5/187—Convection cooling
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
    • F01D5/12—Blades
    • F01D5/14—Form or construction
    • F01D5/18—Hollow blades, i.e. blades with cooling or heating channels or cavities; Heating, heat-insulating or cooling means on blades
    • F01D5/186—Film cooling
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2240/00—Components
    • F05D2240/10—Stators
    • F05D2240/12—Fluid guiding means, e.g. vanes
    • F05D2240/122—Fluid guiding means, e.g. vanes related to the trailing edge of a stator vane
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2240/00—Components
    • F05D2240/20—Rotors
    • F05D2240/30—Characteristics of rotor blades, i.e. of any element transforming dynamic fluid energy to or from rotational energy and being attached to a rotor
    • F05D2240/304—Characteristics of rotor blades, i.e. of any element transforming dynamic fluid energy to or from rotational energy and being attached to a rotor related to the trailing edge of a rotor blade
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2260/00—Function
    • F05D2260/20—Heat transfer, e.g. cooling
    • F05D2260/202—Heat transfer, e.g. cooling by film cooling
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2260/00—Function
    • F05D2260/20—Heat transfer, e.g. cooling
    • F05D2260/221—Improvement of heat transfer
    • Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
    • Y02—TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
    • Y02T—CLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION
    • Y02T50/00—Aeronautics or air transport
    • Y02T50/60—Efficient propulsion technologies, e.g. for aircraft

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  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)

Description

本発明は、タービンバケット翼形部を冷却するための冷却システムに関し、具体的には、半径方向冷却通路を使用して翼形部を対流冷却しかつ翼形部の後縁領域を対流及びフィルム冷却するようになった冷却システムに関する。   The present invention relates to a cooling system for cooling a turbine bucket airfoil, and more particularly to convectively cool an airfoil using a radial cooling passage and convection and film the trailing edge region of the airfoil. The present invention relates to a cooling system that has come to cool.

長年にわたり、ガスタービンは、出力及びエンジン効率を向上させるために入口燃焼温度を高くする傾向にあった。ガス温度が高くなるにつれて、より高い金属温度のバケット翼形部は、大きなクリープ損傷を示す。クリープ損傷は、クリープ破断が発生してタービン流路構成要素を破損するまでに悪化する。さらに、燃焼温度が上昇した場合に、高温燃焼ガスが翼形部の先端の方向に向かって半径方向外向きに遠心力を受けるので、入口温度プロフィールは、翼形部先端においてより高温になる傾向をもつ。このことにより、翼形部のより高いスパン位置はクリープ損傷の影響を一層受け易い。また、翼形部の形状に起因して、翼形部がより高い温度になった時、後縁は、酸化、クリープ及び低サイクル疲労割れを含む次第に大きな損傷を示すようになる。   Over the years, gas turbines have tended to increase the inlet combustion temperature to improve power and engine efficiency. As the gas temperature increases, the higher metal temperature bucket airfoil exhibits greater creep damage. Creep damage is exacerbated before creep rupture occurs and damages turbine flow path components. In addition, the inlet temperature profile tends to be higher at the tip of the airfoil as the combustion temperature rises because the hot combustion gases are subject to centrifugal force radially outward toward the tip of the airfoil. It has. This makes the higher span position of the airfoil more susceptible to creep damage. Also, due to the shape of the airfoil, when the airfoil is at a higher temperature, the trailing edge will show progressively greater damage including oxidation, creep and low cycle fatigue cracking.

従来のバケット設計では、電解加工（ＥＣＭ）法を用いて、ほぼ半径方向にバケットの翼形部を貫通する冷却通路を形成してきた。具体的には、冷却孔は成形チューブ電解加工（ＳＴＥＭ）法を用いて形成され、この加工法により、円形ガイドチューブと同一形状である通路を電気化学的に侵食する円形ガイドチューブを用いて孔を「穿孔する」。ガイドチューブ穿孔加工は、翼形部の長さ全体にわたって延びる。一般的に、バケットのロータホイールへのダブテール取付け部の底面において別個の穿孔パスを開始して、該穿孔パスを翼形部ＳＴＥＭ穿孔加工によって侵食した通路と合流させる。従って、穿孔通路を通ってバケットのダブテールから翼形部の先端に送られる冷却空気は、翼形部を対流冷却する。しかしながら、穿孔した孔の長さが増大すると、成形チューブは、その穿孔方向がずれる傾向になる。また、翼形部は、後縁が空気力学的効率のために非常に薄くなるように成形されている。その結果、ＳＴＥＭ穿孔した孔は、後縁から一定距離の範囲内に穿孔することができるのみで、必然的に後縁の大部分が依然として相対的に冷却されないままの状態になるおそれがある。   In conventional bucket designs, electrolytic machining (ECM) methods have been used to form cooling passages that penetrate the bucket airfoil substantially radially. Specifically, the cooling hole is formed by using a formed tube electrolytic processing (STEM) method, and by this processing method, a hole is formed using a circular guide tube that electrochemically erodes a passage having the same shape as the circular guide tube. "Perforate". The guide tube drilling process extends over the entire length of the airfoil. In general, a separate drilling path is initiated at the bottom of the dovetail attachment to the rotor wheel of the bucket to merge the drilling path with the passage eroded by the airfoil STEM drilling. Thus, cooling air that is directed through the perforation passage from the dovetail of the bucket to the tip of the airfoil convects the airfoil. However, as the length of the drilled hole increases, the molded tube tends to shift its drilling direction. The airfoil is also shaped so that the trailing edge is very thin for aerodynamic efficiency. As a result, STEM drilled holes can only be drilled within a certain distance from the trailing edge, which inevitably leaves the majority of the trailing edge still relatively uncooled.

従って、バケット翼形部の後縁領域をより効果的に冷却するシステムに対する必要性が存在する。   Accordingly, there is a need for a system that more effectively cools the trailing edge region of the bucket airfoil.

本発明の好ましい実施形態では、タービンバケット用の冷却システムを提供し、本システムは、翼形部と、基部と、翼形部及び基部間に接合されたプラットフォームとを含む。翼形部は、その半径方向内側端部において冷却媒体の供給源に連通状態になるように該翼形部に沿ってほぼ半径方向に延びる複数の通路を有し、それによって、冷却媒体が通路に沿って翼形部の先端に向かってほぼ半径方向外向きに流れるときに該翼形部を対流冷却するのが好ましい。所定数の通路には、翼形部先端に隣接した、使用済み冷却媒体をタービンの高温ガス通路内に流すための翼形部出口開口部が形成される。複数の通路の少なくとも１つの残りの通路は、翼形部の先端に隣接して、該翼形部の後縁領域に沿ってほぼ半径方向内向きに延びて後縁領域を対流冷却するようになった冷却通路と連通状態になっている。通路は、翼形部の１つの側面に沿いかつ該翼形部の全長の中間に位置する出口孔で終端して、後縁をフィルム冷却するようになっている。   In a preferred embodiment of the present invention, a cooling system for a turbine bucket is provided, the system including an airfoil, a base, and a platform joined between the airfoil and the base. The airfoil has a plurality of passages extending generally radially along the airfoil such that the airfoil is in communication with a source of cooling medium at a radially inner end thereof, whereby the cooling medium passes through the airfoil. Preferably, the airfoil is convectively cooled as it flows substantially radially outward toward the tip of the airfoil. The predetermined number of passages are formed with airfoil exit openings adjacent the airfoil tips for flowing the spent cooling medium into the hot gas passages of the turbine. At least one remaining passage of the plurality of passages extends substantially radially inward along the trailing edge region of the airfoil adjacent the tip of the airfoil to convectively cool the trailing edge region. It is in communication with the cooling passage. The passage is adapted to film cool the trailing edge, terminating at an exit hole located along one side of the airfoil and midway through the length of the airfoil.

本発明の別の好ましい実施形態では、翼形部と、基部と、翼形部及び基部間に接合されたプラットフォームとを有するタービンバケットを冷却する方法を提供する。本方法は、それに沿ってほぼ半径方向に翼形部先端に向かって延びる複数の通路を有する翼形部を準備する段階と、通路に沿って翼形部先端に向かって冷却媒体をほぼ半径方向外向きに流して翼形部を対流冷却するようにする段階と、使用済み冷却媒体がタービンの高温ガス通路内に流れるのを可能にする、所定数の通路の出口開口部を翼形部先端に隣接して設ける段階と、翼形部の先端に隣接する位置において、複数の通路の残りの通路の少なくとも１つからの冷却媒体を翼形部の後縁領域に沿ってほぼ半径方向内向きに延びる冷却通路と連通状態にして、後縁領域を対流冷却するようにする段階と、翼形部の１つの側面に沿って該翼形部の全長の中間に位置する出口孔で通路を終端させて、後縁をフィルム冷却するようにする段階とを含む。   In another preferred embodiment of the present invention, a method of cooling a turbine bucket having an airfoil, a base, and a platform joined between the airfoil and the base is provided. The method includes providing an airfoil having a plurality of passages extending along the passageway substantially radially toward the airfoil tip, and substantially cooling the cooling medium along the passageway toward the airfoil tip. The airfoil tip has a predetermined number of passage outlet openings that flow outward to allow convective cooling of the airfoil and to allow spent coolant to flow into the hot gas passage of the turbine. A cooling medium from at least one of the remaining passages of the plurality of passages substantially radially inward along the trailing edge region of the airfoil at a location adjacent to the tip of the airfoil The trailing edge region is convectively cooled and communicated with a cooling passage extending to the end of the air passage along one side of the airfoil and terminated with an outlet hole located midway through the airfoil And allowing the trailing edge to cool the film. .

次に図面、特に図１を参照すると、全体を符号１０で表した従来型のタービンバケットを示しており、タービンバケット１０は、翼形部１２と、基部１４と、翼形部１２及び基部１４間に接合されたプラットフォーム１６とを含む。タービンバケット１０は、タービンの高温ガス流路（図示せず）内の高温ガスをシールするための、エンゼルウィングシール１８と先端シュラウド２０とを含む。さらに、翼形部１０は、基部１４からプラットフォーム１６及び翼形部１２を貫通し先端シュラウド２０を貫通してガス流路内に開口する複数のほぼ半径方向に延びる通路２２を含む。一般的に、空気はほぼ半径方向外向きに流れて翼形部先端から高温ガス通路内に流出するので、空気などの冷却媒体をこれら通路に供給して翼形部１２を対流冷却するようにする。図１に示すように、相対的に冷却されない状態のままである後縁２６に隣接する領域２４が存在する。既述のように、通路２２は、後縁領域２４の構造的健全性を維持しながら、後縁２６に十分に近接させて「穿孔する」ことができない。   Referring now to the drawings, and in particular to FIG. 1, a conventional turbine bucket, generally designated 10, is shown which includes an airfoil 12, a base 14, an airfoil 12 and a base 14. And a platform 16 joined therebetween. The turbine bucket 10 includes an angel wing seal 18 and a tip shroud 20 for sealing hot gas in a hot gas flow path (not shown) of the turbine. In addition, the airfoil 10 includes a plurality of generally radially extending passages 22 that extend from the base 14 through the platform 16 and the airfoil 12, through the tip shroud 20, and open into the gas flow path. Generally, air flows substantially radially outward and flows out of the airfoil tip into the hot gas passages so that a cooling medium such as air is supplied to these passages to convectively cool the airfoil 12. To do. As shown in FIG. 1, there is a region 24 adjacent to the trailing edge 26 that remains relatively uncooled. As already mentioned, the passage 22 cannot “pierce” sufficiently close to the trailing edge 26 while maintaining the structural integrity of the trailing edge region 24.

図２及び図３において、また本発明の好ましい実施例によると、例えば空気などの冷却媒体の少なくとも一部分は、特に図１の従来技術によっては解決されない翼形部の後縁領域におけるより大きな冷却を実現するように経路変更される。本発明の冷却システムでは、冷却空気は、後縁に隣接する通路に沿ってほぼ半径方向内向き方向に流れ、翼形部の側面に沿ってフィルムとして吐出される。通路及び出口孔は、後縁を対流冷却すること及び後縁領域に沿って断熱フィルムを形成することの二重の機能を果たす。   2 and 3, and according to a preferred embodiment of the present invention, at least a portion of the cooling medium, such as air, provides greater cooling, particularly in the trailing edge region of the airfoil that is not solved by the prior art of FIG. The route is changed to be realized. In the cooling system of the present invention, cooling air flows in a generally radially inward direction along a passage adjacent to the trailing edge and is discharged as a film along the sides of the airfoil. The passages and outlet holes serve the dual function of convectively cooling the trailing edge and forming a thermal insulation film along the trailing edge region.

図２に示す翼形部３０の好ましい実施例では、複数の通路３２が形成され、翼形部の全長に沿ってほぼ半径方向に延びる。これら通路３２は、前述のＳＴＥＭ穿孔法を用いて穿孔される。所定数の通路３２が、ＳＴＥＭ法で使用する成形チューブの直径の方向ずれ特性を考慮しながら可能な限り後縁３４に近接させて、翼形部の先端３６からほぼ半径方向内向き方向に穿孔される。従って、通路３２の１つ又はそれ以上の通路３８は、可能な限り後縁３４に近接して翼形部に沿ってほぼ半径方向に延びるように形成される。これらの１つ又はそれ以上の通路３８は、翼形部の最も冷却を必要とする臨界領域すなわち後縁領域４０において翼形部の１つの側面、好ましくは正圧側面を横方向に貫通して開口する。通路３８は、断面が円形であるのが好ましい。しかしながら、通路が翼形部の側面を突破するとき、こうして形成した孔４２は、ほぼ楕円形輪郭を有する。応力場に存在する孔４２の楕円形輪郭により、発生する応力集中の影響を軽減する。   In the preferred embodiment of the airfoil 30 shown in FIG. 2, a plurality of passages 32 are formed and extend substantially radially along the entire length of the airfoil. These passages 32 are drilled using the STEM drilling method described above. A predetermined number of passages 32 are drilled in a generally radially inward direction from the airfoil tip 36 as close as possible to the trailing edge 34 while taking into account the directional deviation characteristics of the diameter of the forming tube used in the STEM method. Is done. Accordingly, one or more passages 38 of the passage 32 are formed to extend generally radially along the airfoil as close to the trailing edge 34 as possible. These one or more passageways 38 extend laterally through one side, preferably the pressure side, of the airfoil in the critical or trailing edge region 40 where the airfoil needs to be most cooled. Open. The passage 38 is preferably circular in cross section. However, when the passage breaks through the side surface of the airfoil, the hole 42 thus formed has a generally elliptical profile. The elliptical contour of the hole 42 present in the stress field reduces the effect of the stress concentration that occurs.

１つ又はそれ以上の通路３８内に冷却媒体流を供給するために、翼形部先端の半径方向内側にかつ該翼形部先端を貫いて陥凹部が形成される。陥凹部４４は、１つ又はそれ以上の半径方向通路３２と連通状態になっている。図示した実施形態では、通路３２の２つの通路４６が、翼形部先端に隣接して陥凹部４４と連通状態になっているが、１つ又はそれ以上の通路４６を設けることができることが分かるであろう。また、陥凹部４４は、翼形部先端３６に隣接した、通路３８の１つ又はそれ以上と連通状態になったプレナムを形成することも分かるであろう。従って、冷却空気は、１つ又はそれ以上の半径方向通路４６に沿ってほぼ半径方向外向きに流れ、そこで流れはプレナム４４に流入しかつ流れ方向を逆方向にされて、通路３８に沿って半径方向内向きに流れるようになる。通路３８内の流れは、翼形部の正圧側面に沿って孔４２を通って流出する。通路３８内の冷却空気のほぼ半径方向内向きのこの流れの結果として、翼形部の後縁領域は対流冷却される。また、後縁領域の正圧側面に沿って出口孔４２を通して冷却空気の薄いフィルムが供給されて後縁領域をフィルム冷却する。   A recess is formed radially inwardly through and through the airfoil tip to provide a coolant flow into the one or more passageways 38. The recess 44 is in communication with one or more radial passages 32. In the illustrated embodiment, the two passages 46 of the passage 32 are in communication with the recess 44 adjacent to the airfoil tip, but it will be appreciated that one or more passages 46 may be provided. Will. It will also be appreciated that the recess 44 forms a plenum adjacent to the airfoil tip 36 and in communication with one or more of the passageways 38. Accordingly, the cooling air flows approximately radially outward along one or more radial passages 46 where the flow enters the plenum 44 and is reversed in direction of flow along the passage 38. Flows inward in the radial direction. The flow in the passage 38 flows out through the hole 42 along the pressure side of the airfoil. As a result of this generally radially inward flow of cooling air in the passage 38, the trailing edge region of the airfoil is convectively cooled. A thin film of cooling air is supplied along the pressure side surface of the trailing edge region through the outlet hole 42 to cool the trailing edge region.

図４を参照すると、翼形部３０の先端には、シュラウドカバー４８が設けられる。シュラウドカバーは、シュラウドシール５０と共に、タービンの固定シュラウド（図示せず）に沿って溝を形成するカッタ歯とを含む。図４に示すように、通路３２は、シュラウドカバー４８を貫通する出口開口部５４を有し、これにより翼形部の大部分を対流冷却する冷却媒体を高温ガス通路内に放出する。しかしながら、プレナム４４は、ろう付け、溶接又は他の方法で翼形部の先端に固定することができるカバー又はキャッププラグ５６で覆われる。従って、プレナム４４はカバーキャップ５６で閉鎖されて、１つ又はそれ以上の所定の通路４４を通ってほぼ半径方向外向きに流れる冷却空気が、流れ方向を逆方向にされてほぼ半径方向内向き方向に流れて、翼形部の後縁領域を対流冷却しかつ出口孔４２を通って流出した時に後縁領域をフィルム冷却するのを可能にする。   Referring to FIG. 4, a shroud cover 48 is provided at the tip of the airfoil 30. The shroud cover includes, together with the shroud seal 50, cutter teeth that form grooves along a stationary shroud (not shown) of the turbine. As shown in FIG. 4, the passage 32 has an outlet opening 54 that extends through the shroud cover 48, thereby releasing a cooling medium into the hot gas passage that convectively cools most of the airfoil. However, the plenum 44 is covered with a cover or cap plug 56 that can be brazed, welded or otherwise secured to the tip of the airfoil. Accordingly, the plenum 44 is closed with a cover cap 56 so that cooling air flowing substantially radially outward through one or more predetermined passages 44 is reversed in the flow direction and generally radially inward. Flowing in the direction allows convective cooling of the trailing edge region of the airfoil and film cooling of the trailing edge region as it exits through the outlet hole 42.

図３を参照すると、図２におけるのと同じ部分に対して、数字１を前に付けた同様の参照符号を使用する。この実施形態では、プレナム１４４は、翼形部の先端内に非常に深く延ばされ又は形成される。プレナム１６０の底面１６２は、翼形部先端に半径方向内側に間隙を置いて配置されかつ後縁に向かって傾斜している。従って、延長したプレナム１６０は、ＳＴＥＭ穿孔法により図２の実施形態におけるよりも翼形部内で深い範囲まで通路１３８を形成することを可能にする。このようにして、出口孔１４２は、図２の実施形態における孔４２の位置から半径方向内側の位置で翼形部の正圧側面を貫通する。   Referring to FIG. 3, like reference numerals preceded by the numeral 1 are used for the same parts as in FIG. In this embodiment, the plenum 144 is extended or formed very deeply within the tip of the airfoil. The bottom surface 162 of the plenum 160 is disposed radially inward from the airfoil tip and is inclined toward the trailing edge. Thus, the extended plenum 160 allows passage 138 to be formed to a greater extent within the airfoil by STEM drilling than in the embodiment of FIG. In this way, the outlet hole 142 penetrates the pressure side of the airfoil at a position radially inward from the position of the hole 42 in the embodiment of FIG.

現在最も実用的かつ好ましい実施形態であると考えられるものに関して本発明を説明してきたが、本発明は、開示した実施形態に限定されるものではなく、また、特許請求の範囲に記載された符号は、理解容易のためであってなんら発明の技術的範囲を実施例に限縮するものではない。   Although the present invention has been described with respect to what is presently considered to be the most practical and preferred embodiments, the present invention is not limited to the disclosed embodiments and is not limited to the reference numerals recited in the claims. These are for easy understanding, and do not limit the technical scope of the invention to the embodiments.

破線で示した様々な冷却通路を有する従来技術の冷却システムを示す典型的なタービンバケットの側面図。1 is a side view of an exemplary turbine bucket showing a prior art cooling system having various cooling passages indicated by dashed lines. 本発明の好ましい態様による冷却システムを組込んだ翼形部の部分断面図。1 is a partial cross-sectional view of an airfoil incorporating a cooling system according to a preferred embodiment of the present invention. 本発明の別の実施形態を示す、図２と同様の図。The figure similar to FIG. 2 which shows another embodiment of this invention. 翼形部先端及び翼形部のカバーを示す平面図。The top view which shows the airfoil part front-end | tip and the cover of an airfoil part.

符号の説明Explanation of symbols

１０ タービンバケット
１４ 基部
１６ プラットフォーム
３０ 翼形部
３２、４６ 通路
３４ 後縁
３６ 翼形部先端
３８ 後縁に隣接する通路
４０ 後縁領域
４２ 出口孔
４４ クロスオーバプレナム
５６ カバー 10 turbine bucket 14 base 16 platform 30 airfoil 32, 46 passage 34 trailing edge 36 airfoil tip 38 passage adjacent to trailing edge 40 trailing edge region 42 outlet hole 44 crossover plenum 56 cover

Claims (11)

タービンバケット用の冷却システムであって、
翼形部（３０、１３０）と、基部（１４）と、前記翼形部及び基部間に接合されたプラットフォーム（１６）とを含み、
前記翼形部が、該翼形部に沿ってほぼ半径方向に延びる複数の通路（３２、１３２）であって、その各々がその半径方向内側端部において冷却媒体の供給源に連通しているとともに冷却媒体がそれに沿って該翼形部の先端に向かってほぼ半径方向外向きに流れて該翼形部を対流冷却する、複数の通路（３２、１３２）を有し、
前記複数の通路のうちの所定数が、使用済み冷却媒体をタービンの高温ガス通路内に流すための出口開口部（５４）を前記翼形部先端の近傍に有しており、
前記複数の通路の少なくとも１つの残りの通路（４６、１４６）が、前記翼形部の先端の近傍で、該翼形部の後縁の領域（４０、１４０）に沿ってほぼ半径方向内向きに延びて該後縁領域を対流冷却する冷却通路（３８、１３８）と連通しており、
前記冷却通路（３８、１３８）が、前記翼形部の一方の側面に沿いかつ該翼形部の全長の中間に位置する出口孔（４２、１４２）で終端して、該一方の翼形部側面に沿って前記後縁をフィルム冷却する、
冷却システム。 A cooling system for a turbine bucket,
An airfoil (30, 130), a base (14), and a platform (16) joined between the airfoil and the base,
The airfoil is a plurality of passages (32, 132) extending substantially radially along the airfoil, each communicating with a source of cooling medium at its radially inner end. And a plurality of passageways (32, 132) along which a cooling medium flows substantially radially outward toward the tip of the airfoil to convectively cool the airfoil,
A predetermined number of the plurality of passages has an outlet opening (54) in the vicinity of the airfoil tip for flowing a used cooling medium into the hot gas passage of the turbine;
At least one remaining passage (46, 146) of the plurality of passages is generally radially inward along a trailing edge region (40, 140) near the tip of the airfoil. In communication with cooling passages (38, 138) for convectively cooling the trailing edge region,
The cooling passage (38, 138) terminates in an outlet hole (42, 142) located along one side of the airfoil and in the middle of the entire length of the airfoil, Film cooling the trailing edge along the side,
Cooling system. 前記出口孔（４２、１４２）が、前記翼形部の正圧側面を貫通して開口する、請求項１記載のシステム。 The system of claim 1, wherein the outlet hole (42, 142) opens through a pressure side of the airfoil. 前記出口孔（４２）が、前記正圧側面に沿って楕円形輪郭を有する、請求項２記載のシステム。 The system of claim 2, wherein the outlet hole (42) has an elliptical profile along the pressure side. 前記冷却通路（３８、１３８）が、前記少なくとも１つの残りの通路よりも前記後縁（３４）に近接した経路に沿って延びる、請求項１乃至請求項３のいずれか１項記載のシステム。 The system according to any of the preceding claims, wherein the cooling passage (38, 138) extends along a path closer to the trailing edge (34) than the at least one remaining passage. 前記複数の通路（３２、１３２）の残りの通路（４６、１４６）が、前記翼形部の先端の近傍で、該翼形部の後縁領域に沿ってほぼ半径方向内向きに延びて該翼形部の後縁領域を対流冷却する少なくとも２つの冷却通路（３８、１３８）と連通しており、前記少なくとも２つの冷却通路（３８、１３８）の各々が、前記翼形部の一方の側面に沿いかつ該翼形部の全長の中間に位置する出口孔（４２、１４２）で終端して、該一方の翼形部側面に沿って前記後縁をフィルム冷却する、請求項１乃至請求項４のいずれか１項記載のシステム。 The remaining passages (46, 146) of the plurality of passages (32, 132) extend substantially radially inward along the trailing edge region of the airfoil , near the tip of the airfoil. In communication with at least two cooling passages (38, 138) for convectively cooling the trailing edge region of the airfoil, each of the at least two cooling passages (38, 138) being on one side of the airfoil. A film cooling of the trailing edge along the side surface of the one airfoil, terminating at an outlet hole (42, 142) located along the middle of the airfoil and along the length of the airfoil. 5. The system according to any one of 4. 前記出口孔（４２、１４２）が、前記翼形部の全長に沿って異なる半径方向位置に位置する、請求項５記載のシステム。 The system of claim 5, wherein the outlet holes (42, 142) are located at different radial locations along the entire length of the airfoil. 前記翼形部（３０）が、前記翼形部先端の近傍に、前記少なくとも１つの残りの通路（４６、１４６）の出口と前記冷却通路（３８、１３８）への入口とに連通したクロスオーバプレナム（４４、１４４）を含む、請求項１乃至請求項６のいずれか１項記載のシステム。 A crossover in which the airfoil (30) communicates with the outlet of the at least one remaining passage (46, 146) and the inlet to the cooling passage (38, 138) in the vicinity of the tip of the airfoil. The system according to any one of the preceding claims, comprising a plenum (44, 144). 前記プレナム（４４、１４４）が、前記翼形部の先端を貫いて開放しており、カバー（５６）が前記翼形部の先端において該プレナムを閉鎖している、請求項７記載のシステム。 The system of claim 7, wherein the plenum (44, 144) is open through the tip of the airfoil and a cover (56) closes the plenum at the tip of the airfoil. 前記プレナム（４４）が、前記翼形部先端から半径方向内側に間隙を置いて配置された底面（１６２）を有する、請求項８記載のシステム。 The system of claim 8, wherein the plenum (44) has a bottom surface (162) spaced radially inward from the airfoil tip. 前記底面（１６２）が後縁に向かって傾斜している、請求項９記載のシステム。 The system of claim 9, wherein the bottom surface (162) is inclined toward a trailing edge. 前記複数の通路（３２、１３２）の各々が成形チューブ電解加工（ＳＴＥＭ）穿孔法を用いて穿孔されたものである、請求項１乃至請求項１０のいずれか１項記載のシステム。
The system according to any one of the preceding claims, wherein each of the plurality of passageways (32, 132) is perforated using a molded tube electrochemical machining (STEM) perforation method.
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